Internal cooling circuit for blade outer air seal formed of laminate

Patent metadata

Field	Value
Publication number	US-10822985-B2
Application number	US-201816116140-A
Country	US
Kind code	B2
Filing date	Aug 29, 2018
Priority date	Aug 29, 2018
Publication date	Nov 3, 2020
Grant date	Nov 3, 2020

Abstract

A method of forming a gas turbine engine component includes the steps of (a) forming an intermediate portion, (b) forming cooling circuit structure into at least an outer layer of the intermediate portion, (c) providing an outer layer over the formed cooling circuits to close off the cooling circuits, such that there are laminate on both a radially inner and a radially outer side of the cooling circuits, and (d) forming an inlet and an outlet to the cooling circuits through the outer layer. A gas turbine engine is also disclosed.

First claim

The invention claimed is:

1. A method of forming a gas turbine engine component comprising the steps of:
(a) forming an intermediate portion;
(b) forming cooling circuit structure into at least an outer layer of said intermediate portion;
(c) providing an outer layer over said formed cooling circuits to close off said cooling circuits, such that there are laminate on both a radially inner and a radially outer side of said cooling circuits; and
(d) forming an inlet and an outlet to said cooling circuits through said outer layer;
wherein said intermediate portion includes a plurality of laminate;
wherein at least one inner layer is also added inward of said intermediate portion in step (c); and
wherein an overwrap layer provides both said inner and outer layer.

2. The method as set forth in claim 1 , wherein said component is a blade outer air seal.

3. The method as set forth in claim 1 , wherein step (d) is performed by an ultrasonic machine.

4. The method as set forth in claim 3 , wherein there are a plurality of separate cooling circuits formed within said intermediate layer.

5. The method as set forth in claim 4 , wherein there is an individual inlet and an individual outlet for each of said plurality of cooling circuits.

6. The method as set forth in claim 5 , wherein said cooling circuit has a non-rectangular shape.

7. The method as set forth in claim 1 , wherein said intermediate portion and said layers are formed of a ceramic matrix composite.

8. The method as set forth in claim 1 , wherein step (b) is performed by an ultrasonic machine.

9. The method as set forth in claim 8 , wherein step (d) is performed by an ultrasonic machine.

10. The method as set forth in claim 1 , wherein there are a plurality of separate cooling circuits formed within said intermediate layer.

11. The method as set forth in claim 1 , wherein said cooling circuit has a non-rectangular shape.

12. A gas turbine engine comprising:
a compressor section and a turbine section, said turbine section including at least one rotor and at least one blade extending radially outwardly from said rotor to a radially outer tip;
a blade outer air seal assembly positioned radially outwardly of said radially outer tip of said blade, said blade outer air seal having forward and aft hooks, and said forward and aft hooks being supported on forward and aft seal hooks of an attachment;
said blade outer air seal formed of a plurality of laminate layered with a central web formed of a plurality of laminate members including an inner reinforcement member, and an outer overwrap that wraps around said inner reinforcement member; and
said blade outer air seal forward of a plurality of laminate layers, with internal cooling circuits formed in at least one of said layers, with at least one other layer radially outward of said at least one layer, and closing off said internal cooling circuit.

13. The gas turbine engine as set forth in claim 12 , wherein an inlet and an outlet is provided through said at least one other layer radially outward of said at least one layer, for each of said internal cooling circuits.

14. The gas turbine engine as set forth in claim 13 , wherein there being an intermediate portion forming said internal cooling circuits, and said at least one other layer radially outward of said at least one layer being positioned outwardly of said intermediate portion.

15. A blade outer air seal comprising:
a blade outer air seal assembly having forward and aft hooks;
said blade outer air seal formed of a plurality of laminate layered with a central web formed of a plurality of laminate members including an inner reinforcement member, and an outer overwrap that wraps around said inner reinforcement member; and
said blade outer air seal forward of a plurality of laminate layers, with internal cooling circuits formed in at least one of said layers, with at least one other layer radially outward of said at least one layer, and closing off said internal cooling circuit.

16. The blade outer air seal as set forth in claim 15 , wherein an inlet and an outlet is provided through said at least one other layer radially outward of said at least one layer, for each of said internal cooling circuits.

17. The blade outer air seal as set forth in claim 16 , wherein there being an intermediate portion forming said internal cooling circuits, and said at least one other layer radially outward of said at least one layer being positioned outwardly of said intermediate portion.

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Clark Thomas E
• Smiddy Winston Gregory
• Whitney Daniel J
• Barker William M

Classifications

• F05D2250/11

  triangular · CPC title

• F05D2250/13

  trapezoidal · CPC title

• F01D11/24

  by selectively cooling-heating stator or rotor components · CPC title

• F05D2260/204

  by the use of microcircuits · CPC title

• F05D2240/11

  Shroud seal segments · CPC title